This invention relates to a flow structure for a gas turbine, in particular for an aircraft engine.
Gas turbines such as aircraft engines usually have multiple compressors, multiple turbines and a combustion chamber. The multiple compressors usually include a low-pressure compressor and a high-pressure compressor; the multiple turbines usually include a high-pressure turbine and a low-pressure turbine. The gas turbine has flow passing through it axially; the low-pressure compressor is situated upstream from the high-pressure compressor, and the high-pressure turbine is situated upstream from the low-pressure turbine. The flow goes from the low-pressure compressor into the high-pressure compressor through a transitional channel between these two compressors. Likewise, such a transitional channel is positioned between the high-pressure turbine and the low-pressure turbine.
It is already known from the prior art that supporting ribs spaced a distance apart may be provided in such transitional channels between two compressors or two turbines in the circumferential direction of the transitional channel. The supporting ribs are used in providing, for example, oil lines and sensors and for receiving forces, which is why the supporting ribs are designed to be relatively thick. There are known supporting ribs in the prior art which are designed to carry but not guide the flow. In addition, there are also known supporting ribs which have a suction side and a pressure side and therefore also assume the function of flow guidance. In particular in the case of supporting ribs that deflect the flow, there is the risk of flow separation and the risk of development of secondary flow, which ultimately results in high flow losses, because of the small height ratio and the relatively great thickness of the supporting ribs in the case of supporting ribs that deflect flow. In addition, this also results in a poor oncoming flow quality for the components situated downstream from the transitional channel in the direction of flow.
Against this background, the present invention is based on the problem of creating a novel flow structure for a gas turbine.
This problem is solved by a flow structure for a gas turbine of the present invention. According to this invention, at least one guide vane and/or guide rib is positioned between two adjacent supporting ribs arranged a distance apart in the circumferential direction of the transitional channel whereby the flow outlet edge of the guide rib or each guide rib runs upstream from the flow outlet edges of the supporting ribs.
With the help of the inventive flow structure, an optimized flow is obtained within the transitional channel, thereby minimizing the risk of the development of flow separation and secondary flow. In addition, there is optimum oncoming flow for a component downstream from the flow arrangement and/or the transitional channel. Flow losses are avoided and the efficiency of the gas turbine is optimized.
The flow inlet edge of the guide rib or each guide rib positioned between two adjacent supporting ribs preferably runs slightly upstream from the flow entrance edges of the supporting ribs and in addition a channel wall bordering the transitional channel on the inside radially and/or a channel wall bordering the transitional channel on the outside radially is also constricted inward in the area of the flow outlet edge of the guide rib or each guide rib positioned between two neighboring supporting ribs.